Ejections of satellites at low spin speed are currently problematic because transverse spring ejection torque impulse creates a precession error inversely proportional to the spacecraft spin speed. The firing of thrusters to correct nutation adds to the problem. The present invention permits the reduction of allowable spin speed at separation by enabling an autonomous correction of the sun angle postejection, then maintaining this sun angle and nearly eliminating attitude walk in the transverse direction.
U.S. Pat. No. 4,370,716, to Amieux, discloses an active nutation control system for a space vehicle. The system disclosed controls nutation with thrusters and maintains attitude by bookkeeping the transverse angular impulse and restricting firings to those that correct both nutation and precession. The use of a sun sensor as the nutation sensor is mentioned in passing in column 7, line 64, but direct precession measurement is not suggested.
While the prior techniques function with a certain degree of efficiency, none discloses the advantages of the satellite spin vector control with sun sensor of the present invention as is hereinafter more fully described.